Notes

Summary:

Summary: Tests were performed in the Göttingen (Germany) and Guidonia (Italy) supersonic tunnels in order to determine the aerodynamic characteristics of projectiles of various shapes. The Mach numbers ranged from about 1.3 to 3.2 for the Göttingen tests and from 1.44 to 2.66 for the Guidonia tests. The results show that increasing the relative length of the nose causes the drag coefficient to decrease and the center of pressure to move forward. For a given length, the nose having minimum drag has a curved profile; the curvature is greatest at the tip and decreases to a very small value toward the rear of the nose, where the shape becomes approximately conical. As the Mach number increases, the drag coefficient decreases and the center of pressure moves toward the tail. For the higher Mach numbers the variation of the drag coefficient and the movement of the center of pressure are small. Existing aerodynamic theory gives values of the aerodynamic characteristics close to those determined experimentally for small flow deviations.

"NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were previously held under a security status but are now unclassified. Some of these reports were not technically edited. All have been reproduced without change in order to expedite general distribution."

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SUPERSONIC-TURUEL TESTS OF PROJECTILES IN
GERMANY AND ITALY
By Antonio Ferri

Langley Memorial Aeronautical
Langley Field, Va.

WASHINGTON

Laboratory

NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of
advance research results to an authorized group requiring them for the war effort. They were pre-
Sviously held under a security status but are now unclassified. Some of these reports were not tech-
nically edited. All have been reproduced without change in order to expedite general distribution.
L 152
DOCUMENTS DEPARTMENT

Tests were performed in the Gottingen (Germany) and
Guidonia (Ittly) supersonic turnels in order to determine
the aerodynam.ic characteristics of Tprojectiles of various
shares. The I':ach numbers rar;-ed from. atout 1.5 to 5.2
for the Gottingen tests and fromrr 1.44 to 2.66 for the
Guidonia tests. The results show that increasing the
relative length of the nose causes the drs-e coefficient
to decrease In-l the center of pressure to move forward.
For a given length, the nose haviai;i ,.ninnmum dreg has a
curved profile; the curvature is cr'rac'st at the tip and
decreases to a very smsil value covward the rear of the
nose, where the shape oeccnomes soproc'--imately conical. As
the Mach nu-mber increases, the ir-g coefficient decreases
and the canter of pressure mDves toward the tail. For
the hi-her ,erch numbers the vari;tLion of the drag coef-
ficient and the movement of the center of pressure are
small. Existing aerod:ynmamic theory gives values of the
aerodynamic characteristics clost to chcsm determined
experimentally for small flow. deviations.

IlTROiDUCT IC!

Rcssarch on nroj:ctilas was ctart.:d at rppr-jximately
the sanehe Gotc ineTn Lr.borat.ory in Gjrmany
and the C-uidrni. Lboratcry in Italy. 'Th3 original data
were brought to the United Ststes in C10 and were tabu-
lated Prd analyz-_d at the Lan:,lay L-.eorial A.-ron-iutical
Laboratory of th Fa'stional Advisory Coj.r.itt-;' for Aero-
nautics.

The eim of the 3ermen res.; rch 117,s tc determine the
variation of Rerodrynamic characteristics with Wach number
for various fundamental geomettriz.l sliapes for projectiles.

2 CONFIDEIITIAL :iACA ACHR 11i. L5:1O0

Force tests of small models O.953 inch in diameter were
conducted in a supersonic tunnel having : a test section
approximately 2.36 by 2.84 inches. The shape of che
models was systematically varied to determine the effects
on the aerod:.-namic characteristics of:

(1) Nose profile F.hoe for a tvpicel fineness ratio

(2) Nose length for noses with circular-arc profiles

(53) Smll taper of the tail of the projectile

The program carried out at Che Guidonia Laboratory
had as its aim the de'elomnent of rn ontimum -hTnpe for a
1.812-inch-caliber antitar.k projectile. Eequse of the
relatively l-rge size of the test model, it was possible
to obtain pressure-distribution studies as well as precise
aerc.dynsjnic data for cojrmpri-.ion wiLh results derived by
existing, theory.

The theories presented in references 1 to 6 for
sharp-nose projectil-s at zero en,.',lu of attsc-K and the
theories of references 7 and 3 for bodies of revolution
at an anicle of yaw were used to caDnute the theorjtical
characteristics of the various conical noses for com-
narison with the e:xnerinmental results. The aerodn',ric
theory of rinir:mum-drs" projectiles rresentsI in refer-
ences 3 snd 5 to 7 was us.ed as a guide in th. design of
the various nose shapes cf ttn projectiles tested.

SYMBOLS

The symbols used for defininT the aerod..nemic coef-
ficients and the --;onmetric characteristics of the pro-
jectiles are riven in figure 1.

n fineness retio of nose (Zn/,
E angle between intersection of ten!lnt on nose of
orojectile End generezrix of' vliinder

COI:FIDENTIAL

CO"FTIDTEN.TTAL

CiACA ACR :No. L5H08

E P_'ERI:.']ETS AT GOT'ITN-E;:

Nind-Tunnel mad Experimental Methods

Ex:erlments were carried out by Gerran technicians
of the Aeroiynemische Versuchsansbalt (AVA) in the small
sa. ersonic tunnel Pt the Gottingen Laboratory. The
lroyout of the wind tunnel is shown in figure 2. The
tunnel has a rectcngul--r section with a throat about 2.36
by 2.8u4 inches. As indicated in figure 2, a semiopen
throat arrangement was used. The side walls of the tunnel
were strpi:jht and parallel, but the jet was not restrained
by too &nd bottom walls. It has been found that this
;'.renc-rr.*ent rrEkes it possible to obtain reliable aero-
dyr.snmic data at MPch numbers onl:, slightly greater
t~1-. 1.0 (reference 9) end haet the choking condition
which would exist if the jet were com!ietely restrained
does !not occur. Some trouble was encountered during the
tests because of ccndJnza9tion i'-ienoomens, in saite of the
fact t.hat the numi'.dity of the &nolent .ir had ace-n reduced
to a low vilue by -r-eliminsry drying.

Th6 tests consisted of the i.ieasurement of lift, drag,
and oitchin- moment with a semi autoira-ic balance. Each
model ''es castedi hrao0.;'h an nP1le-of-attE.ck range from 90
to -30 rndj a .-..c0l nu.mbc r range from 1.5 to 5.2. For the
lonr er models, it wa;s not possible to ,erforr the tests
at the low v,0lozitie Lec-use th- front shock wave
reflecting frcrr the jet bomundrries interfered with the
flow on the rear 1fce of the projectile.

Test Models

The models each had a di:reter of 0.O95 inch and were
supported by a sting attached to the rear face. The
dimensions of tnr,- sting and the tare system adopted are
not known.

The models can be separated into three distinct
groups to determine the effect of:

Nose shope.- Four projectiles having over-all fineness
ratios n of 5.0 with nose fineness ratios of 2.5 modelss
1, 2, 3, and fIg- 5) wer,! tested to determine the
effect of the nose shape. because the models were small,
they I.ad no tail tLper or bourrslet ring. All the noses

COC 1F IDETTI AL

CO-!FIDEfTIAL

NACA ACR io. L5HO8

were of circular profile with the radii r varying
from 6.5d to co (conical nose). For the nose with
6.5d radius (model 1, fiF. 5), the end of the nose was
tangent to the cylinder at their juncture but, for the
other models, the profile of the nose terminated with
its tangent inclined at an angle with respect to the
generatrix of the cylinder.

Length of the nose.- Five models having nose lengths
varying from C.5d to 6.5d (models 5, 6, 1, 7, and 8,
fig. t) were tested to determine the effect of the length
of the nose. The noses of all the models had circular
profiles tangent to the cylinder forming the body of the
projectile.

Tail taper.- Three modeis derived from model 1 and
having three different tail tapers (9:odels 9, 10, and 11,
fig. 5) were tested to determine the effect of tail taper.

Results

Wind-tunnel tests.- Tht results of the experiments
at thL-. 'cttii'ren L6torato.ry are given in figures 6 to 21.
Figures & to 11 show the results of tests to ^determine
the effect of the nose. sht'pe. In figures 6 to 9 the
variation of the aerodynmiic coefficients for models 1,
2, 3, and 4 is shown for sev-.ral anrwles of attack and a
rarne of .sch n-umbers. Fiures 10 and 11 show the aero-
dynamic coefficients c.f each ,;rojectile at equal .;ach
num-bers as a function of th, ratio d/r. As shown in
figures 10 and 11, the mini:r.um: drec' coefficient was
obtained for a nose intermaidlate to the noses for which
r = co and r = 12.5d. The difL f.rnces in Lhe minimum
drag coefficients were not large.

The results of the tests to determine the effect of
the length of the ncse are given in figures 12 to 17.
Increasing the fineness ratio of the nose nT caused
the drqa c':efficient to drop noticeably (fir. 16). The
slope dCD/dnn decreased as n1i increased. The center
of pressure of th3e projectiles moved toward the nose as
the finene33 re tic of c!.: n.os, inca'.,sc ei (fig. 17).

The results of the tests to evaluate the effect of
the tail taper ere given in figures 18 to 21. The vari-
ation of the aerodynamic coefficients with M.cch number is

CONFIDENTIAL

COI.I DENT I AL

A IlACA, ACR No. L5HOS

shown in figures 18 to 20, and data for comparison of the
ta' ered models with model 1 are shown in figure 21. 'Th-
dra. of the projectile was lowest for the longest tapered
tail, especially at the lower Mach numbers (fig. 21).
The differences in drag coefficients were not large. The
other aercdyT:amic characteristics were not appreciably
altered.

The drag coefficient for a given projectile shape
decreased as the i.;ach number was increased. This effect
was most pronounced at the lower su:.ersonic velocities;
the vsriaticn -t ",Mch numbers of the order of 5.0 was
slight. The position of the center of pressure did not
change appreciably with anqle of attack. In the lower
supersonic velocity range the center of pressure moved
toward the rear of the projectile ,s the speed was
increased but tended to eporoach a fixed location at the
higher TMch numbers.

Firing tests.- Actual firing tests were performed
to verify tihe experimental values, and the following
results were obt.-ain .d:

ICD
od Firing tests Tunnel tests
(a f) (a : )
1 O.Lo 0.29
1 2 :.6 .27

.I .28

Those results were for a !.'ich nu.-nber of 2.2.

The projectiles used in the firing tests had an
.anle of attack of nearly 5 end : bouri-'elet r.n. Drag
coefficients obtained from firing tests at angles of
attack near 30 gavu drag coefficients that were equive-
lent to a wind-tunnel angle of attack of about 7.50.
The differences therefore cann-ot be entirely due to the
presence of the bourrelet nor to the error in angle of
attack. The differences may probably be attributed in
part to the difference of surf,-ce finish between the
tunnel model and the fired projectiles, the rotation of
the fired projectile about its axis, and the difference
in Reynolds number for the firing and the tunnel tests.

CO1F FIDZETI AL

COIFIDEITTIAL

NACA ACR i'o. L5H8O3

EXPERITRNTS AT ,UIDONIA

Wind-Tunnel and Exoer'imentF.l M!ethcds

Projectiles with nine different nose shares :.ere
tested in the closed-throat high-speed turnel at 3uidonia
(reference 10) at rPech numbers ra.-n'Iing froir 1.14 to 2.u6.
The system. of the rpPrtially open streak vyes not used
because it required a larger nrnouiit of ncwer' snd there-
fore limited the maximum velocity. The Lest section was
large enough not to require special attentionn to preventt
choking of the air streEmrr with the rrodel in the tunnel
when the Miach ni.nLmocr was greater then 1.1,. The nozzles
ware of recarg.ulcr s-ection of th- two-aimensional type
with s minimum section 1..7*- by 15.74 inches.

The forces on eech rn-od1l wer-e determined by use of
a three-comnorent b.lrnce (rsierence 10). The model was
attached to the balances by sting or thle projectile
axiom on rr fce th rr fe nroiectile. The sting,
althou..;h of .11 : :ll er, affected the exe-rimencel
results sc:ne":iht irc.e :t incr:-eas3e'. t.ai 'Presrure on the
rear ;rce o' th'e -cjec(il]e. It w.s n-ece-ss .r, therefore,
to ra rn- acr.rj.e t.e? r.ea.-t.Lre.rient cb' russending the
mod'l on 2 23'reJ vtruc tinttached to the cid3 of the
procjectile,

found: 0.87 for a Mach number of 1.85 and 0.36 for a iach
number of 2.13. Thece earlier results, however, are
questionable becFuse the effect of the support strut,
which increases the pressure at the rear somewhat, was
ne -lected.

Test Projectiles

A 1.812-inch-celiber sntitani projectile was used for
the body of the projectile in th2e Quidonia tssts, the
details of .vhich are shown in figure 22. Nine different
nose shapes (fig. 23) were tested with this body in order
to dete-rrine the nose for minimum drag.

Conical rises of varying i injness ratio (projec-
tiles 1, 2, 3, and 4 of fie. 25) vere t-sted first in
ordC-r to estimate the importance of fineness ratio. The
simple conical nos,- form: was chosen- to permit ccmprrison
with existin- thaoretic'-l data for conicEl noses.

Both the a-proxinimte projectile th-.ories (refer-
ences i. and 5) end the comol.-x but' more exact theory of
Perr-,ri (references 6, p rt I, and 7) show th.t, in
order to minimize. the drag of the projectile, it is
necessary to conce-ntrat.e th. ores.2ur-e at th- vertex of
the nose and then to carry out tlho iLost rapid expq~osion
possible. :.Vth t!;is criterion as e. uide, five noses of
fineness ratio np- = 2.0 were d3sig.ned (projectiles 5,
6, 7, 8, and 9 of fig. 2;). (See t'ble I.) Two of the
noses tested, projectiles 3 and 9, were blunt-endeJ but
were otherwise sLmiler to the nose of projectile 5. The
nose of projectile 5 anprosches that theoretically derived
by Ferrari (reference o). A nose exactly corresponding
with the ootimum nose described by the theory of Ferrari
was about to be tested when the tc sts v'ere suspended.

Results

The results of the experiments at Guidonia are shown
in figures 2!1 to 35. The variation of CD and f/d
with fineness retio cf the nose at a Msch number of 2.06
anls an angle of attack of 0 is shown in figure 24; the
varitior. of CD and f/d with iM'.ch number for the nose
h9.'ir;. a finu.n;css ratio of 2.0 (~rojectile 3) at an angle
of attack of 00 is shown in figure 25. The drac coef-icient

COTFIDENTTAL

NACA ACR No. L5H08

for a given projectile shape decree as'd as the Ma;ch number
increased. The position of the center of pressure did
not change appreciably with angle of attack. The values
obtained from firing tests for two noses having fineness
ratios of 2.0 and 2.5 at a iLach number of 2.14 are also
shown in figure 24. The argle of attack. in th- firing
tests varied between 20 and 30.

The dreg vslu,-s 'ivrTn by the firing tests are som.-
whqt higher then those determined in the wind tunnel. The
difference can probably be attributed to the f'ct th.-t the
models in the tunnel were perfectly finished but the fired
projectiles had rough machine finish. The fired pro-
jectile also had a r'otstionEl rmctior that v.:as not ,rpro-
ducesd in the tunn-;l tests snd thct indcoubtcdly ltcr1d
th- phenomenon of the boundary lay-Jr.

The fcllowin- tables show calculated values of the
initial shock-vave ringle and the pressure coefficient on
the nose for nm-ojectiles 1 qanid 5 at Much n'rLrber of 2.06.
These quaprtities i\'er-e calculated by the method of refer-
ences 7 and 3. Exrerimertal values of these quantities
are shown for comparison.

(2) At the higher angles of yaw appreciable
differences exist between the theoretical end experimental
pressure coefficients, particularly if the nose is short.

(3) Pressure on the rear faces of the projectile is
only slightly affected by the nose s}laoe but is epreciably
decreased with an increase of nrigl of attack.

Figure 30 shows the vnri.qtion of the a-rodynamic
coefficients with engle of attack for projectile 3, and
figure 31 gives the or.rs-ure distributionss for this
rrojectile at a Mach number of 2.0S. The following

CO .F I DEN TI AL

CONFIDENTIAL

----

-ohenoie "..8 -."? close

NAC.I -CR No. L5HOb CONFIDaETIAL .11

table comoares the drag cofficier.ts for a = 0 and
the center-of-reossure posiulons for the various projec-
tiles for = 2.06:

It will b- observed thAt projectiles 5 and 9 had
the lowest drag. These three .shr1es c.re closer to the
optimum profile predcct3d cheo'etic'!.Iy than any cf the
other noses tested. The pressure-distribution diagrams
(fig. 51) and the flow photographs (firs. 54 and 55) show
that, when the front part of tlh projectile is flat as
for projectiles 6 and 9, a rormin-l shock wa'e occurs and
the pressure at the nose nLproacles The street tctal
pressure in value. The shock iwave is detached fro.n the
projectile. Irmitediately behind the blur-nc face. cf the
nose a. raid expansion occurs, and the pressures a short
distance frcm the nos,: be.coi.e lower then for the conical
nose. These lower pressures act over a ruel.tively largs
part of the frontal pres :f the projectile; consequently,
a lower drrg coefficient is obtained for the blunt nose
than for the conical nose. The pressure on the rear face
of the projectile is about the same for both types of
nose. Tne lift at the same anile of attack for the blunb
noses is -lightly greater than for the conical ncses, and
the center of pressure is therefore further f.orwav.p
These differences are very small, however. It riay be
mentioned that the blunc type of nose is more practical
than the shaer-pointed nosc from the standpcints of con-
struction end rd.iiin t.;nanc,.

COi.TCIUSI .TS

Tests ware prforme. in th'2 Gottingen (Germany) and
Guidonia (Italy) sunrrsoniic tL'nnels in oroer to deternminc
the aerodynmsmic chsrr.ct-ristic.c of -.rojectjle of various

'OJT'ID DENI IAL

N;.CA LCRi No. L5HOS

shapes. The follo'.:ing conclusions era besed on the results
of both the German and the Italian exopri:nents:

1, The fineness ratio of the nose is of primary impor-
tance in determining the aerodynamic characteristics of
supersonic nrojectiles. As the fineness ratio increases,
the drag coefficient decreases and the center of pressure
moves forward.

2, The dra- coefficient for s given projectile shape
decreases as the MIach number is increased. This effect
is most pronounced at Lhe lov:r sui:ersonic velocities;
the variation at MNcn numbe'-rs of the order of 3.0 is slight.

3. The position of tne center of' -ressure does not
change appreciably with Engle of attack. In the lower
supersonic velocity rwA,.e te center of pressure moves
toward the rear of the projectile as ihe speed is increased
but tends to E.pproch fixecd location at the higher Mach
numbers,

The pr-,ssure cn tni rear face of the projectile
varies apcreciably with an:le o.f Itt&aci but is only
slightly affected by *the formn of the inmse.

5. For a given fineness ratio th- optimum nose
profile has e relrLively blunt en, which is fired to
the cylindrical p&rt of the p.rojrctii-;. The theortical
criterions for thu desisln of the optinu:n. nose profile
have been v-..ified.

6. The existing ferod'nnpmic theory.r for the calcu-
lation of the pCezsure dis ctbution ab-out 'riojt:Liles is
adequately precise for snall flow deveitions.

7, The ed'iition of a .small taper Lc the tail of the
projectile diminishes the drag slirh-tly, particularly at
the I~-rlrr 'Mch numbers, vichout .ltii:ring thm other aero--
dynamic characteristic:.

9. Ferri, Antonio: Completed Tabulation in the United
States of Tests of 24 Airfoils at Eigh Mach Numbers
(Derived from Interrupt-d Work at Guidonia, Italy,
in the 1.51- by 1.74-Foot High-Speed Tunnel). NACA
ACF No. L5E21, 1945.

Figure 6.- Lift, drag, and moment coel'icients and position
of center of pressure as functions of Mach number for
various angles of attack. Model 1. (G6ttingen).

Fig. 6

O

I I
-- eq79 ___

OC

1 3

CONFIDENTIAL
I I

NACA ACR No. L5H08

2.0

1.6

2.6

2.4-

20
/.

C.

-("9-

NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS

) I I
-t 1 1

CONFIDFNTI AL
I I I I

/.69 26
A/Vrch umb-er,- A^,

Figure 6.- Concluded.

QI
QlJ

( L1

I

r

=

Fig. 6 Cone.

(.

3.0 0..4

NACA ACR No. L5H08

/. 9 2.2? 2.6
/1ofch number, /Wo

Figure 7.- Lift, drag, and moment coefficients and position
of center of pressure as functions of Mach number for
various angles of attack. Model 2. (G6ttingen).

Fig. 7

NACA ACR No. L5H08

U
.qI

Q)
0
k

Ob

LI
o
0

Ik
Li

2.0

/.6

/.2

6

.4

J.6

3.2

2.6

2.4

2. 0

Figure 7.- Concluded.

/. ?.? 2.6
,,W^lchn number, /1\

Fig. 7 Cone.

NACA ACR No. L5H08

_de.92-

___-- --..--.---- .---- -- -- -- --

CONFIDENTIAL
'.0 .4 /.< Z. Z ."6 3.0 J.
A Cac/ number, /-f

Figure 8.- Lift, drag, and moment coefficients and position
of center of pressure as functions of Mach number for
various angles of attack. Model 3. (G6ttingen)

Fig. 8

NACA ACR No. L5H08

2.O

1.6

1.2

4

Fig. 8 Cone.

1 /.4 /.6 2.2 2.6
/aych- nufmbeLt, /o

I I
3.0 d3.,4

Figure 8.- Concluded.

CONFIDENTIAL

S-. -. --
0

,3 3_

-o---5 -
U

NATIONAL ADVISORY
COMMITTEE FO AERONAUTICS
S i l l

CONFIDENTIAL
I I 1

2.4

2.0

=

" '

NACA ACR No. L5H08

S i I I
CONFIDENTIAL

(de9y -- -

9
6

0

NATIONAL ADVISORY
COMMITTEE FO1 AERONAUTICS

cX
(ode9-

C I

CONFIDENTIAL
I I I

Mac h number, /Io

Figure 9.- Lift, drag, and moment coefficients and position
of center of pressure as functions of Mach number for
various angles of attack. Model 4. (Gottingen)

2.6 t.0 0.4-

/.0 /.4-

/.6 2.2

Fig. 9

NACA ACR No. L5H08

/.6d .Z 2.6
/faoch number, Mo

- Figure 9.- Concluded.

.9,
u

1'c
U

i,
qj
U

LI

"--
Qi
U

Ri
k
Qi;

.1^
(5 Q
Pr

NATIONAL ADVISORY
CONNITTEE FR AEIOrUTiKS

2.4

. O

3.4

Fig. 9 Cone.

NACA ACR No. L5HO8

0
(j
'1}

8

?3

0 .05
o///

./5

Figure 10.- Lift, drag, and moment coefficients and position
of center of pressure as functions of the ratio of the
diameter of the body of the projectile to the radius of
the nose of the projectile for various angles of attack.
Mo a 2.0. (Gottingen)

Figure II.- Lift, drag, and moment coefficients and position
of center of pressure as functions of the ratio of the
diameter of the body of the projectile to the radius of
the nose of the projectile for various angles of attack.
Mo = 3.0. (G6ttingen)

NACA ACR No. L5H08

S.8

joe

0 -
,^ ^2:
$ t -

3.2

S28

24

2O0

Model 4

Fig. 11 Cone.

I I I
0o I Mode/l
(d5e -g -
6190)~5

Wodel/ 3

I-I- 1l l ; I i l l

CONFIDENTIAL NATIONAL ADVISORY
COMMITTEE FOI AERONAUTICS

.05
d/

./0

./5

Figure 11.- Concluded.

"

Model 2Jl
- ^ =V__I_

r

NACA ACR No. L5H08

A~c~~Ch n/6 et;Y

Figure 12.- Lift, drag, and moment coefficients and position
of center of pressure as functions of Mach number for
various angles of attack. Model 5. (Gottingen)

Fig. 12

NACA ACR No. L5H08

/8 22 26
^7ac^ /-7^1??e -sq; /e

Figure 12.- Concluded.

U

Qj

U'
q)

A`

/2

.B

.4

z4

20

/2

NATIONAL ADVISORY
CONNITTEE FO AERONAUTICS

Fig. 12 Cone.

NACA ACR No. L5H08

YI
1 '

/e?-I Z.- -A-e)
>^/^>/>^/}er 7

Figure 13.- Lift, drag, and moment coefficients and position
of center of pressure as functions of Mach number for
various angles of attack. Model 6. (Gottingen)

9

/ 0-0-

CONFIDENTIAL
L; I I

Fig. 13

tY

5 u

NACA ACR No. L5H08

28.

24

20

/4!

J.0

Figure 13.- Concluded.

.en,
'3,
qj0

"c
005C

CONFIDENTIAL

Fig. 13 Cone.

/8 2-' 2.6
zv;*c1 num6en /47-,

34

NACA ACR No. L5H08

NATIONAL ADVISORY
CUMMIIIEE FOR AERONAUIICS

/.6 2.2 26
/Aac/h number, /-,,

Figure 14. Lift, drag, and moment coefficients and position
of center of pressure as functions of Mach number for
various angles of attack. Model 7. (Gottingen)

9
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'3

CONFIDENTIAL

Fig. 14

3.0

3.4

NACA ACR No. L5H08

3.6

0

kcb

3.2

26

2.4

2.

Fig. 14 Cone.

3.0

3.4

Figure 14.- Concluded.

CONFIDENTIAL

9

,-,----

c _C

NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS

CONFIDENTIAL
S i_____

Mcrch nur;Y7ber,

NACA ACR No. L5H08

kj .6

CONFIDENTIAL

9

NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS

6 _

3

CONFIDENTIAL
I I I

/8 d2 2.6
/^icych number, ,9

J.0

3.4

Figure 15.- Lift, drag, and moment coefficients and position
of center of pressure as functions of Mach number for
various angles of attack. Model 8. (Gottingen)